1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a large span air cooled turbine rotor blade for an industrial gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
Latter stages of turbine blades do not require film cooling air, but do require internal convection cooling in order to control the metal temperature to within acceptable levels in order to provide for a long service life. FIG. 1 shows one such prior art turbine blade with radial cooling passages from the root to the blade tip in which convectional cooling occurs. The FIG. 1 blade includes three radial cooling channels 11-13 each having skewed trip strips or turbulators formed along the walls that function to enhance the heat transfer efficiency of the cooling channel. FIG. 2 shows a cross section top view of the blade of FIG. 1 with the radial cooling channels and trip strips along the walls.
In the radial cooling channels with trip strips of FIGS. 1 and 2, as the cooling air flows through the skewed trip strips, the leading edge of the trip strip trips the thermal boundary layer of the cooling air which results in a higher local heat transfer coefficient and thus an increase in the airfoil cooling performance. A normal flow of cooling air over a flat surface would produce a boundary layer between the flat surface and the moving cooling air flow. The boundary layer acts as a baffle zone. Tripping the boundary layer using the trip strips increases the heat transfer rate. FIG. 3 shows the prior art skewed trip strips along the radial passage with the leading edge at a lower spanwise height than the trailing edge of the same trip strip. FIG. 4 shows a cross section view through one of the trip strips in FIG. 3 along the line A-A with the cooling air flow paths over the trip strips. A result of this boundary layer tripping is that vortices are generated and propagate along the trip strips from the leading edge to the trailing edge. As these vortices propagate along the full length of the trip strip, the boundary layer becomes progressively more disturbed or thick, and therefore the tripping of the boundary layer becomes progressively less effective. The result of this boundary layer growth is a significantly reduced heat transfer effect. Also, for a large channel height cooling air passage typical for latter stage industrial engine turbine blades, the vortex occurs near the inner wall of the airfoil. A majority of the cooling air flow still remains in the middle of the radial passage away from the hot wall surface that requires the convection cooling.